The present invention relates to attitude control of a spacecraft and, in particular, to ensuring accurate payload pointing performance.
For a spacecraft to perform its mission generally requires precise orientation control of the spacecraft payload relative to its target. For example, the antenna of a geosynchronous communications satellite must generally be pointed at a fixed location on the earth. Spacecraft that use a zero-momentum control system require an on-board attitude determination system to accurately determine the spacecraft orientation in order to control the attitude. Based on the estimated attitude, the control logic applies a torque to the spacecraft to correct the error between the desired and estimated attitude.
A well established method of spacecraft attitude determination relies on an inertial measurement unit (IMU) to measure spacecraft rate. The rate measurement is used to propagate the estimated spacecraft attitude; biases in the IMU gyro outputs will therefore corrupt the attitude estimate and degrade payload pointing performance. In order to improve the accuracy of the propagated attitude estimate, the spacecraft yaw, roll, and pitch angles are measured using a complement of sensors such as an earth sensor, sun sensor, and star tracker. The propagated attitude from the IMU and the measured spacecraft angles are used by an extended Kalman filter (EKF) to calculate errors in the propagated attitude and the gyro biases. The estimation error and bias outputs of the EKF are used to correct the estimated attitude and the IMU rate measurements.
A major shortcoming of the prior-art attitude determination system is the degradation in estimated attitude due to angular distortions between the IMU and the attitude sensor(s) as well as errors in the attitude sensor outputs. As described below, a major component of these error sources occurs at orbit frequency due to thermal distortion between the IMU and the attitude sensor(s) as well as due to changes in the sensor temperatures.
The present invention provides a method for correcting spacecraft pointing errors. According to the method an orbit frequency distortion angle between a rate sensor and at least one attitude sensor is estimated. An estimated attitude of the spacecraft is adjusted to correct for the distortion angle.
The present invention also provides a method for correcting payload pointing errors of a spacecraft. The method includes sensing the spacecraft attitude with an attitude sensor. An inertial measurement unit angular rate is determined. An orbit frequency component of an angle between an inertial measurement unit and attitude sensor is estimated to determine an angle between the payload and a reference frame of the attitude sensor.
The present invention also includes a spacecraft attitude control system that implements the above-described methods.